Gas turbine engine

The present disclosure relates to a gas turbine engine, and more specifically to a gas turbine engine having a nacelle.

A gas turbine engine typically includes a fan and a turbomachine. The turbomachine generally includes a nacelle in which an inlet, one or more compressors, a combustor, and one or more turbines are disposed. The one or more compressors compress air which is then channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the one or more turbines which extract energy from the combustion gases for powering the one or more compressors, as well as for producing useful work to propel an aircraft in flight. The turbomachine is mechanically coupled to the fan for driving the fan during operation.

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C.

The term “turbomachine” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc., each refer to relative speeds within an engine unless otherwise specified. For example, a “low speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high speed turbine” of the engine.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

As used herein, the terms “axial” and “axially” refer to directions and orientations that extend substantially parallel to a centerline of the gas turbine engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the centerline of the gas turbine engine. In addition, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately about the centerline of the gas turbine engine.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

A “third stream” as used herein means a non-primary air stream capable of increasing fluid energy to produce a minority of total propulsion system thrust. The third stream may generally receive inlet air (air from a ducted passage downstream of a primary fan) instead of freestream air (as the primary fan would). A pressure ratio of the third stream may be higher than that of the primary propulsion stream (e.g., a bypass or propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of an airflow through the third stream with a primary propulsion stream or a core air stream, e.g., into a common nozzle.

In certain exemplary embodiments an operating temperature of the airflow through the third stream may be less than a maximum compressor discharge temperature for the engine, and more specifically may be less than 350 degrees Fahrenheit (such as less than 300 degrees Fahrenheit, such as less than 250 degrees Fahrenheit, such as less than 200 degrees Fahrenheit, and at least as great as an ambient temperature). In certain exemplary embodiments these operating temperatures may facilitate heat transfer to or from the airflow through the third stream and a separate fluid stream. Further, in certain exemplary embodiments, the airflow through the third stream may contribute less than 50% of the total engine thrust (and at least, e.g., 2% of the total engine thrust) at a takeoff condition, or more particularly while operating at a rated takeoff power at sea level, static flight speed, 86 degree Fahrenheit ambient temperature operating conditions.

Furthermore in certain exemplary embodiments, aspects of the airflow through the third stream (e.g., airstream, mixing, or exhaust properties), and thereby the aforementioned exemplary percent contribution to total thrust, may passively adjust during engine operation or be modified purposefully through use of engine control features (such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features) to adjust or optimize overall system performance across a broad range of potential operating conditions.

Generally, a gas turbine engine is attached to a wing of an aircraft via a pylon fairing. The space between the wing and the gas turbine engine adjacent to the pylon fairing is referred to as a gully area. Airflow through the gully area is a three-sided channel that passes over a lower surface of the wing, a side surface of the pylon fairing, and an upper surface of the gas turbine engine. The limited space of the gully area causes the airflow to accelerate, and the lower surface of the wing, the side surface of the pylon fairing, and the upper surface of the gas turbine engine create drag on the airflow, also known as scrubbing. The drag is typically worse on an inboard side of the pylon fairing.

The inventor of the present disclosure discovered that, while the drag can be an issue for ducted fans in which the airflow through the gully area is freestream flow, the drag is worse for open fan structures due to the airflow being supercharged through the unducted fan prior to passing through the gully area. The airflow pressurized by the unducted fan may be accelerated by a throat of the gully area, i.e., a section with a minimum area, then may further accelerate at the expanding flow area downstream of the throat. The unducted fan increases the Mach number of the airflow passing through the gully area, which may result in higher drag at the lower surface of the wing, the side surface of the pylon fairing, and the upper surface of the gas turbine engine. Supersonic flow and/or excessive drag can form in the gully area. The higher Mach number of the airflow through the gully area can result in a strong shock and a wave drag penalty that decreases efficiency. The inventors recognized that increasing the radial distance between the lower surface of the wing and the upper surface of the gas turbine engine would increase the area at the throat, reducing the acceleration at the throat of the gully area, in turn reducing the Mach number of the airflow through the gully area. This may result in reduced drag at the lower surface of the wing, the side surface of the pylon fairing, and the upper surface of the gas turbine engine, reduce the shock at the gully area, and improve efficiency of the aircraft.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,shows a schematic cross-sectional side view of a gas turbine engineaccording to one or more embodiments of the present disclosure. Particularly,provides an engine having a rotor assembly with a single stage of unducted rotor blades. In such a manner, the rotor assembly may be referred to herein as an “unducted fan,” or the entire gas turbine enginemay be referred to as an “unducted engine.” In addition, the engine ofincludes a third stream extending from the compressor section to a rotor assembly flowpath over the turbomachine, as will be explained in more detail below.

For reference, the gas turbine enginedefines an axial direction A, a radial direction R, and a circumferential direction C. In the below explanation, radial direction R is defined as a vertical direction of the installed gas turbine engineon a wing(). Moreover, the gas turbine enginedefines an axial centerline or longitudinal axisthat extends along the axial direction A. In general, the axial direction A extends parallel to the longitudinal axis, the radial direction R extends outward from and inward to the longitudinal axisin a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred sixty degrees (360°) around the longitudinal axis. The gas turbine engineextends between a forward endand an aft end, e.g., along the axial direction A.

The gas turbine engineincludes a turbomachineand a rotor assembly, also referred to a fan section, positioned upstream thereof. Generally, the turbomachineincludes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. For example, as shown in, the turbomachineincludes a core cowlthat defines an annular core inlet. The core cowlfurther encloses at least in part a low pressure system and a high pressure system. For example, the core cowldepicted encloses and supports at least in part a booster or low pressure (“LP”) compressorfor pressurizing the air that enters the turbomachinethrough core inlet. A high pressure (“HP”), multi-stage, axial-flow compressorreceives pressurized air from the LP compressorand further increases the pressure of the air. The pressurized air stream flows downstream to a combustorof the combustion section where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air.

It will be appreciated that as used herein, the terms “high/low speed” and “high/low pressure” are used with respect to the high pressure/high speed system and low pressure/low speed system interchangeably. Further, it will be appreciated that the terms “high” and “low” are used in this same context to distinguish the two systems, and are not meant to imply any absolute speed and/or pressure values.

The high energy combustion products flow from the combustordownstream to a high pressure turbine. The high pressure turbinedrives the high pressure compressorthrough a high pressure shaft. In this respect, the high pressure turbineis drivingly coupled with the high pressure compressor. The high energy combustion products then flow to a low pressure turbine. The low pressure turbinedrives the low pressure compressorand components of the fan sectionthrough a low pressure shaft. In this respect, the low pressure turbineis drivingly coupled with the low pressure compressorand components of the fan section. According to one or more embodiments, the LP shaftis coaxial with the HP shaft. After driving each of the turbines,, the combustion products exit the turbomachinethrough a turbomachine exhaust nozzle.

Accordingly, the turbomachinedefines a working gas flowpath or core ductthat extends between the core inletand the turbomachine exhaust nozzle. The core ductis an annular duct positioned generally inward of the core cowlalong the radial direction R. The core duct(e.g., the working gas flowpath through the turbomachine) may be referred to as a second stream.

The fan sectionincludes a fan, which is the primary fan according to one or more embodiments. In the embodiments shown in, the fanis an open rotor or unducted fan. As depicted, the fanincludes an array of fan blades(only one shown in). The fan bladesare rotatable, e.g., about the longitudinal axis. As noted above, the fanis drivingly coupled with the low pressure turbinevia the LP shaft. For the embodiments shown in, the fanis coupled with the LP shaftvia a speed reduction gearbox, e.g., in an indirect-drive or geared-drive configuration.

Moreover, the fan bladescan be arranged in equal spacing around the longitudinal axis. Each bladehas a root and a tip and a span defined therebetween. Further, each fan bladedefines a fan blade tip radius Ralong the radial direction R from the longitudinal axisto the tip, and a hub radius (or inner radius) Ralong the radial direction R from the longitudinal axisto the base.

Moreover, each bladedefines a central blade axis. In one or more embodiments, each bladeof the fanis rotatable about their respective central blades axes, e.g., in unison with one another. One or more actuatorsare provided to facilitate such rotation and therefore may be used to change a pitch of the bladesabout their respective central blades axes.

The fan sectionfurther includes a fan guide vane arraythat includes fan guide vanes(only one shown in) disposed around the longitudinal axis. In the embodiments shown in, the fan guide vanesare not rotatable about the longitudinal axis. Each fan guide vanehas a root and a tip and a span defined therebetween. The fan guide vanesmay be unshrouded as shown inor, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanesalong the radial direction R or attached to the fan guide vanes.

Each fan guide vanedefines a central blade axis. For the embodiments shown in, each fan guide vaneof the fan guide vane arrayis rotatable about their respective central blades axes, e.g., in unison with one another. One or more actuatorsare provided to facilitate such rotation and therefore may be used to change a pitch of the fan guide vaneabout their respective central blades axes. However, in other embodiments, each fan guide vanemay be fixed or unable to be pitched about its central blade axis. The fan guide vanesare mounted to a nacelle.

As shown in, in addition to the fan, which is unducted, a ducted fanmay be included aft of the fan, such that the gas turbine engineincludes both a ducted and an unducted fan which both serve to generate thrust through the movement of air without passage through at least a portion of the turbomachine(e.g., without passage through the HP compressorand combustion section for the embodiment depicted). The ducted fan is rotatable at about the same axis as the fan blade. The ducted fanis, for the embodiments depicted, driven by the low pressure turbine(e.g. coupled to the LP shaft). In the embodiments depicted, as noted above, the fanmay be referred to as the primary fan, and the ducted fanmay be referred to as a secondary fan. It will be appreciated that these terms “primary” and “secondary” are terms of convenience, and do not imply any particular importance, power, or the like.

The ducted fanincludes a plurality of fan blades (not separately labeled in). The fan blades of the ducted fanmay be arranged in equal spacing around the longitudinal axis. Each blade of the ducted fanhas a root and a tip and a span defined therebetween. Further, each fan blade of the ducted fandefines a fan blade tip radius Ralong the radial direction R from the longitudinal axisto the tip, and a hub radius (or inner radius) Ralong the radial direction R from the longitudinal axisto the base.

The nacelleannularly encases at least a portion of the core cowland is generally positioned outward of at least a portion of the core cowlalong the radial direction R. Particularly, a downstream section of the nacelleextends over a forward portion of the core cowlto define a fan flowpath or fan duct. According to one or more embodiments, the fan flowpath or fan ductmay be understood as forming at least a portion of the third stream of the gas turbine engine.

Incoming air may enter through the fan ductthrough a fan duct inletand may exit through a fan exhaust nozzleto produce propulsive thrust. The fan ductis an annular duct positioned generally outward of the core ductalong the radial direction R. The nacelleand the core cowlare connected together and supported by a plurality of substantially radially-extending, circumferentially-spaced stationary struts(only one shown in). The stationary strutsmay each be aerodynamically contoured to direct air flowing thereby. Other struts in addition to the stationary strutsmay be used to connect and support the nacelleand/or core cowl. In one or more embodiments, the fan ductand the core ductmay at least partially co-extend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl. For example, the fan ductand the core ductmay each extend directly from the leading edgeof the core cowland may partially co-extend generally axially on opposite radial sides of the core cowl.

The gas turbine enginealso defines or includes an inlet duct. The inlet ductextends between an engine inletand the core inlet/fan duct inlet. The engine inletis defined generally at the forward end of the nacelleand is positioned between the fanand the fan guide vane arrayalong the axial direction A. The inlet ductis an annular duct that is positioned inward of the nacellealong the radial direction R. Air flowing downstream along the inlet ductis split, not necessarily evenly, into the core ductand the fan ductby a splitter or leading edgeof the core cowl. The inlet ductis wider than the core ductalong the radial direction R. The inlet ductis also wider than the fan ductalong the radial direction R.

The exemplary gas turbine enginedepicted further includes an array of inlet guide vanespositioned in the inlet ductupstream of the ducted fanand downstream of the engine inlet. The array of inlet guide vanesare arranged around the longitudinal axis. In the embodiment depicted, the inlet guide vanesare not rotatable about the longitudinal axis. Each inlet guide vanesdefines a central blade axis (not labeled), and is rotatable about their respective central blade axes, e.g., in unison with one another. One or more actuatorsare provided to facilitate such rotation and therefore may be used to change a pitch of the inlet guide vanesabout their respective central blades axes. However, in other embodiments, each inlet guide vanesmay be fixed or unable to be pitched about its central blade axis.

Further, located downstream of the ducted fanand upstream of the fan duct inlet, the gas turbine enginemay include an array of outlet guide vanes. As with the array of inlet guide vanes, the array of outlet guide vanesmay not be rotatable about the longitudinal axis. However, for the embodiment depicted, unlike the array of inlet guide vanes, the array of outlet guide vanesmay be configured as fixed-pitch outlet guide vanes.

Further, in the embodiment depicted, the fan exhaust nozzleof the fan ductis further configured as a variable geometry exhaust nozzle. In such a manner, the gas turbine engineincludes one or more actuatorsfor modulating the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to vary a total cross-sectional area (e.g., an area of the nozzle in a plane perpendicular to the longitudinal axis) to modulate an amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flowrate, etc. of an airflow through the fan duct). A fixed geometry exhaust nozzle may also be adopted.

The combination of the array of inlet guide vaneslocated upstream of the ducted fan, the array of outlet guide vaneslocated downstream of the ducted fan, and the fan exhaust nozzlemay result in a more efficient generation of third stream thrust during one or more engine operating conditions. Further, by introducing a variability in the geometry of the inlet guide vanesand the fan exhaust nozzle, the gas turbine enginemay be capable of generating more efficient third stream thrust across a relatively wide array of engine operating conditions, including takeoff and climb (where a maximum total engine thrust may be desirable) as well as cruise (where a lesser amount of total engine thrust may be desirable).

Referring still to, in one or more embodiments, air passing through the fan ductmay be relatively cooler (e.g., lower temperature) than one or more fluids utilized in the turbomachine. In this way, one or more heat exchangersmay be positioned in thermal communication with the fan duct. For example, one or more heat exchangersmay be disposed within the fan ductand utilized to cool one or more fluids from the turbomachine with the air passing through the fan duct, as a resource for removing heat from a fluid, e.g., compressor bleed air, oil or fuel.

Although not depicted, the heat exchangermay be an annular heat exchanger extending substantially 360 degrees in the fan duct(e.g., at least 300 degrees, such as at least 330 degrees). In such a manner, the heat exchangermay effectively utilize the air passing through the fan ductto cool one or more systems of the gas turbine engine(e.g., lubrication oil systems, compressor bleed air, electrical components, etc.). The heat exchangeruses the air passing through ductas a heat sink and correspondingly increases the temperature of the air downstream of the heat exchangerand exiting the fan exhaust nozzle.

is a schematic side view of an aircraft in accordance with one or more embodiments of the present disclosure, andis a schematic view of the gas turbine engineattached to the wingshown inin a reference plane perpendicular to an axial direction A/longitudinal axisand including a gully distance g. As shown in, the gas turbine engineextends from a forward endto an aft endin the axial direction A. The gas turbine engineincludes an array of fan bladesdisposed around the longitudinal axis, which are described in detail above with reference to. As shown in, the gas turbinefurther includes fan guide vanesdisposed on a nacelle (or housing)of the gas turbine enginearound the longitudinal axis, which are also described in detail above with reference to. As shown in, an engine inletis positioned between fan bladesand the fan guide vanes, at a forward end of the nacellein the axial direction A. A fan exhaust nozzleis positioned at an aft end of the nacellein the axial direction A. In the embodiment depicted, the fan exhaust nozzleis formed annularly around the longitudinal axisbetween an inner surface of the nacelleand an outer surface of the core cowl. As explained above with respect to, the fan exhaust nozzleis an exhaust for the fan ductthat may form at least a portion of the third stream of the gas turbine engine. Further, a turbomachine exhaust nozzleis positioned at an aft end of core cowlin the axial direction A.

The wingdefines a lower surfaceon a lower side of the wingwith respect to the radial direction R of the gas turbine engine, and an upper surfaceon an upper side of the wingwith respect to the radial direction R of the gas turbine engine. As shown in, the wingfurther defines a leading edgeon a forward side of the wing(the leading edgeis positioned at a stagnation point when the aircraft including the wingis operated in a cruise operating condition) with respect to the axial direction A of the gas turbine engineand the wingfurther defines a trailing edge (not shown) on an aft side of the wingwith respect to the axial direction A of the gas turbine engine. The gas turbine engineis attached to the wingvia a pylon fairing. In the embodiment depicted, the pylon fairingextends between the lower surfaceof the wing, the leading edgeof the wing, or both to an upper surface of the gas turbine engine. More specifically, for the embodiment depicted, the pylon fairingis attached to an upper surface of the nacelle, the core cowl, or both. Although not shown, according to one or more embodiments, the pylon fairingmay extend to the aft endof the gas turbine engine.

It will be appreciated, however, that the configuration depicted inis by way of example only. In other exemplary embodiments of the present disclosure, the pylon fairingmay be oriented in any suitable manner relative to the wingand/or the gas turbine engine. For example, in certain exemplary embodiments, the pylon fairingmay be oriented perpendicularly to the wing, may be oriented perpendicularly to a plane parallel to the ground when installed, etc.

As shown in, the gully distance g is formed between the wingand the gas turbine engine, and more specifically for the embodiment show is defined as the minimum distance between the wingand the gas turbine engine. More specifically, still, for the embodiment depicted the gully distance g is the minimum distance between the gas turbine engine, as measured from the lower surfaceof the wing(e.g., a surface below the leading edge) to a first point pon an upper surface of the nacelledirectly adjacent to the pylon fairingwith respect to the circumferential direction C of the gas turbine engine. The gully distance g is defined within a reference plane, the reference planeextending perpendicular to the axial direction A and positioned at an axial position.

At the corresponding axial positionof the gully distance g, the gas turbine enginedefines a first distance dthat is a distance from the longitudinal axisto an upper outer surface portion of the nacelle(i.e., an upper half of the outer surface of the nacellein a normal operational attitude), and more specifically to a first point pon an outermost surface of the gas turbine engineon an upper side of the gas turbine engineat a position directly axially adjacent to the pylon fairing, as shown in. At the axial positionof the gully distance g, the gas turbine enginefurther defines a second distance dthat is a distance from the longitudinal axisto a lower outer surface portion of the nacelle(i.e., a lower half of the outer surface of the nacellein a normal operational attitude), and more specifically to a second point pon an outermost surface of the gas turbine engineon a lower side of the gas turbine engine, as shown in. The first point p, the longitudinal axis, and the second point pare colinear (i.e., a straight line passes through all three). As shown in, at the axial positionof the gully distance g, the distance measured radially from the longitudinal axisto the outer surface of the gas turbine engineis maximum at the lowermost portion of the gas turbine engine, and continually decreases in the upper direction until the minimum is reached at the uppermost portion of the gas turbine engine, adjacent to the pylon fairing.

As shown in, gully areasare formed between the wingand the gas turbine engineadjacent to the pylon fairing. The fan exhaust nozzle, which may be an exit of a third stream flowpath according to one or more embodiments, narrows from a lower end of the gas turbine enginein the radial direction R to an upper end of the gas turbine enginein the radial direction R, such that the upper portion of the fan exhaust nozzleis significantly narrower than the lower portion of the fan exhaust nozzle. As a result, the first distance dis less than the second distance d.

Conventional turbofan engines extend symmetrically around a longitudinal axis. That is, for any given position in an axial direction, the outer surface of a conventional gas turbine engine is disposed at an equal distance in the radial direction from the longitudinal axis. As explained above, the first distance dis less than the second distance d. For a given gully distance g between the longitudinal axisand the lower surfaceof the wing, decreasing the first distance dresults in an increase in the gully distance g. Thus, it will be appreciated that, by structuring the gas turbine engine such that the first distance dis less than the second distance d, the gully distance g is increased compared with a conventional gas turbine engine in which the radius of the outer surface is annularly constant.

is a schematic cross-sectional aft view of a gas turbine engineattached to a wingaccording to one or more embodiments of the present disclosure, andis an explanatory view of a flat sectionof the fan nacelle to the right of the pylon fairingincompared with a conventional fan nacelle. Specifically,shows a cross-sectional view of the gas turbine engineattached to the wingat a reference planeat an axial positionin which a gully distance g is defined, similar to the reference planeand axial positionof. Description of the portions similar towill be omitted.

The gas turbine engineshown inincludes a flat sectionat an upper portion of the gas turbine enginenear the pylon fairing. In the embodiment depicted, the nacelleincludes symmetrical sections() that are annularly symmetrical around the gas turbine engine, with the fan exhaust nozzlecorrespondingly annularly symmetrical within the symmetrical sectionsof the nacelle. Additionally, the nacellefurther includes curved sections() immediately adjacent to the pylon fairing. The nacellefurther includes, between the curved sectionsand the symmetrical section, flat sections() that are substantially flat.